This invention relates to aircraft structure elements, particularly skin panels and lower wing stringers for high capacity commercial aircraft, made from rolled, extruded or forged products made of an AlCuMg alloy in the treated temper by solution heat treating, quenching and aging, and introducing a compromise between the different required usage properties that is better than is possible with products according to prior art.
The designations of the alloys and metallurgical tempers used below are according to the designations used by the Aluminum Association, and reused in European Standards EN 515 and EN 573 part 3.
Wings of high capacity commercial aircraft comprise an upper part consisting of a skin made of thick plates made from a 7150 alloy in T651 temper, or a 7055 alloy in T7751 temper or a 7449 alloy in T7951 temper, and stringers made from sections of the same alloy, and a lower part composed of a prefabricated skin made of thick plates of 2024 alloy in the T351 temper or 2324 alloy in the T39 temper, and stringers made from sections of the same alloy. The two parts are assembled by spars and ribs.
The 2024 alloy according to the designations of the Aluminum Association or standard EN 573-3 has the following chemical composition (% by weight):
Si less than 0.5, Fe less than 0.5, Cu=3.8-4.9, Mg=1.2-1.8, Mn=0.3-0.9, Cr less than 0.10, Zn less than 0.25, Ti less than 0.15.
Various alternative solutions have been proposed to improve the compromise between the various required properties, particularly mechanical strength and toughness. Boeing has developed the 2034 alloy with the following composition:
Si less than 0.10, Fe less than 0.12, Cu=4.2-4-8, Mg=1.3-1.9, Mn=0.8-1.3, Cr less than 0.05, Zn less than 0.20, Ti less than 0.15, Zr=0.08-0.15.
This alloy is described in patent EP 0031605 (equivalent to U.S. Pat. No. 4,336,075). Compared with the 2024 alloy in the T351 state, it has a higher specific yield strength due to the increased content of manganese and the addition of another anti-recrystallizing agent (Zr), and improved toughness and fatigue resistance.
U.S. Pat. No. 5,652,063 (Alcoa) concerns an aircraft structure element made starting from an alloy with the following composition (% by weight):
Cu=4.85-5.3, Mg=0.51-1.0, Mn=0.4-0.8, Ag=0.2-0.8, Si less than 0.1, Fe less than 0.1, Zr less than 0.25 and Cu/Mg is between 5 and 9.
The yield strength of the sheet metal made from this alloy in the T8 temper is  greater than 77 ksi (531 MPa). This alloy is intended particularly for supersonic aircraft.
U.S. Pat. No. 5,593,516 (Reynolds) relates to an alloy for aeronautical applications containing 2.5 to 5.5% Cu and 0.1 to 2.3% Mg, in which Cu and Mg contents are kept below their solubility limit in aluminum and are related by the following equations:
Cumax=5.59xe2x88x920.91 Mg and Cumin=4.59xe2x88x920.91 Mg
The alloy may also contain Zr less than 0.20%, V less than 0.20%, Mn less than 0.80%, Ti less than 0.05%, Fe less than 0.15%, Si less than 0.10%.
U.S. Pat. Nos. 5,376,192 and 5,512,112 originating from the same initial patent application are applicable to alloys of this type containing 0.1 to 1% of silver. Note that the use of silver in this type of alloy increases the production cost and creates difficulties in recycling manufacturing scrap.
Furthermore, for many years, xe2x80x9cAU6MGTxe2x80x9d type alloys have been known, according to the old alloy designations in France. Patent FR 1379764 filed by Pechiney in 1963 applies to the use of an alloy of this type with composition Cu=5-7, Mg=0.10-0.50, Mn =0.05-0.50, Si less than 0.30, Fe less than 0.50, Ti=0.05-0.25 for the manufacture of compressed gas cylinders.
The Aluminum Association registered the 2001 alloy in 1976, with the following composition:
Cu=5.2-6, Mg=0.20-0.45, Mn=0.15-0.50, Si less than 0.20, Fe less than 0.20, Cr less than 0.10, Ni less than 0.05, Ti less than 0.20, Zr less than 0.05.
To the best knowledge of the inventors, there is no other industrial use of this alloy apart from compressed gas cylinders manufactured by reverse extrusion.
The current trend in commercial aircraft construction is to use an increasing number of very thick products, with structure elements being machined in the body of these parts. For example, for some small aircraft, wing skins are machined from relatively thick plates to enable in-depth machining of wing stringers, although these stringers are usually made from sections or folded plates and are then mechanically fixed to the skin. Integral in-depth machining of the skin-stringer assembly can reduce manufacturing costs, since there are fewer parts and assembly is avoided. Furthermore, the use of an unassembled structure reduces the weight of the assembly.
Therefore it is desirable that, in addition to the properties normally required for aircraft structure elements, namely high mechanical strength, good tolerance to damage, good fatigue resistance and good resistance to the different forms of corrosion, plates need uniform mechanical properties throughout their thickness, in other words their properties should not vary significantly as a function of the thickness, typically between 10 and 120 mm. Furthermore, the more machining is necessary, the more desirable it becomes to maintain good stability under machining, and this is achieved by a low level of internal stresses. It is known that the mechanical properties for a thick plate are more uniform and internal stresses are lower if the plate is less sensitive to quenching.
Finally, aircraft wings, particularly for high capacity aircraft, have a curved wing profile with curvature in the longitudinal and in the transverse directions. This complex shape can be obtained in an autoclave during the aging process by forming on a mold, by applying a partial relative vacuum on the surface of the mold side of the plate, lower than the pressure on the other side. It is essential that this operation is successful to avoid expensive scrapping of parts with high added value, and particularly large parts. The key to success is in the lowest possible springback effect for a given mold shape, since springback is frequently the most difficult factor to be controlled.
The purpose of this invention is to supply aircraft structure elements with properties at least equivalent to the properties of the same elements made from a 2024 alloy in the T351 temper concerning static mechanical properties, toughness, crack propagation rate and resistance to corrosion, by using rolled, extruded or forged products with low residual stresses, low quench sensitivity and good formability during aging.
The purpose of the invention is a structure element, particularly a lower wing element, manufactured from a rolled, extruded or forged product made of an alloy with composition (% by weight):
Cu=4.6-5.3, Mg=0.10-0.50, Mn=0.15-0.45, Si less than 0.10, Fe less than 0.15, Zn less than 0.20, Cr less than 0.10, other elements  less than 0.05 each and less than 0.15 total, the remainder being Al treated by solution heat treating, quenching, controlled tension to more than 1.5% permanent deformation and aging.
This element has at least one of the following properties:
yield strength R0.2 (TL direction) greater than 350 MPa, and preferably greater than 370 MPa,
toughness K1c (L-T direction) greater than 42 MPam
resistance of P type to intercrystalline corrosion according to standard ASTM G110.
Another purpose of the invention is a manufacturing process for a structure element comprising:
a) casting a plate or a billet with the composition mentioned above,
b) homogenization of this plate or billet,
c) hot transformation of this plate by rolling or of this billet by extrusion or forging to obtain a product thicker than 10 mm,
d) quenching of the hot transformed product,
e) solution heat treating of this product, preferably at a temperature of less than 10xc2x0 C. at the incipient melting temperature of the alloy,
f) controlled tension of the product to obtain a permanent deformation of more than 1.5%,
g) aging of the product at a temperature greater than 160xc2x0 C., possibly together with forming,
h) machining of the product formed until the final shape of the structure elements.
If the product is a piece of sheet metal, the entry temperature to hot rolling is preferably less than the solution heat treating temperature by at least 40xc2x0 C., and even better by at least 50xc2x0 C.
The invention is based on the observation that a 2001 type alloy with some changes to composition and an appropriate manufacturing procedure, can have a set of properties making it suitable for use in aircraft structures, and more particularly in the lower wing parts for high capacity commercial aircraft, also with attractive properties in terms of low quench sensitivity, low residual stresses and good forming ability during aging.
The range of the copper content is significantly lower than for the 2001 alloy, while remaining higher than 2024 and 2034 alloys for lower wing skin, to compensate for the influence of the low magnesium content on the mechanical strength. It is preferable to choose a copper content exceeding 4.8%, or even 4.9% or even 5%. The magnesium content is of the same order of magnitude as in the 2001 alloy, and is preferably between 0.20 and 0.40%. The Cu/Mg ratio is thus almost always greater than 10, unlike what is stated in U.S. Pat. No. 5,652,063 that recommends a Cu/Mg ratio of between 5 and 9.
The manganese content is controlled within a relatively narrow range. If it is below 0.15%, there is a risk that the grain size will be too large; if it is above 0.45%, a non-recrystallized structure is obtained which makes it more difficult to control residual stresses. The preferred range is between 0.25 and 0.40%. Note that for the same reason, the alloy does not contain any anti-recrystallizing elements such as vanadium or zirconium, unlike what is stated in patent U.S. Pat. No. 5,593,516.
The iron and silicon contents are kept below 0.15 and 0.10% respectively, and preferably below 0.09 and 0.08% respectively, to give good toughness. The alloy may contain up to 0.2% of zinc, this addition having a positive effect on the mechanical strength without having any negative effect on other properties such as resistance to corrosion.
The transformation procedure includes casting a plate or a billet, heating or homogenization to a temperature close to the incipient melting temperature of the alloy and hot transformation by rolling, extrusion or forging. If rolling is adopted, it may include one pass called a widening pass in the direction perpendicular to the other passes and intended to improve isotropy of the product. The hot transformation temperature is preferably slightly lower than the temperature that would normally be used by an expert in the subject with reference to the solution heat treatment temperature. Thus, for rolling, the entry temperature is preferably at least 40xc2x0 C. or even 50xc2x0 C. below the dissolution temperature, and the exit temperature is 20 to 30xc2x0 C. below the entry temperature.
The product is then solution heat treated as completely as possible, for example at a temperature of 10xc2x0 C. below the incipient melting temperature of the alloy, while avoiding burning. This temperature is between 520 and 535xc2x0 C. The solution heat treatment quality may be checked by differential enthalpic analysis. The product is then quenched, for example by immersion in cold water, to achieve a cooling rate of between 10 and 50xc2x0 C./s. After quenching, the product is stretched until the permanent deformation is at least 1.5% in order to reduce stresses and improve flatness. For the alloy according to the invention, this tension has the effect of improving the yield strength after aging due to a strain hardening effect, such that the temper obtained can be qualified as a T851 temper, as if it were a specific strain hardening pass after quenching. As mentioned above, aging itself can take place at the same time as the curved shape of the lower wing panel is formed. This aging is preferably done at a temperature exceeding 160xc2x0 C. (and even better greater than 170xc2x0 C.) and sufficiently long to reach the peak yield strength, as for a T6 temper. Typically, aging for a time equivalent to aging for 12 to 24 h at a temperature of 173xc2x0 C. is achieved; any timexe2x80x94temperature combination capable of reaching the alloy aging peak can be used.
The resulting metallurgical structure is strongly recrystallized, unlike the structure obtained with 2024 and 2034 alloys, with a recrystallization rate always exceeding 70%, and usually exceeding 90%, over the entire thickness.
Structure elements according to the invention have compromise properties (static mechanical characteristics, toughness, crack propagation rate, corrosion resistance) that make them suitable for use in aeronautical construction, and particularly for making lower wing skin panels. Furthermore, these elements may easily be made by machining and formed during aging. Finally, the alloy used is easily weldable using standard techniques, so that the number of riveted assemblies can be reduced.
In addition, lower wing elements may be produced according to the invention by machining, in which the skin and stringers are obtained by machining the same initial product.